Gas turbine engine mid turbine frame bearing support

ABSTRACT

A gas turbine engine includes a fan and a compressor section that is fluidly connected to the fan. The compressor includes a high pressure compressor and a low pressure compressor. A combustor is fluidly connected to the compressor section. A turbine section is fluidly connected to the combustor. The turbine section includes a high pressure turbine coupled to the high pressure compressor via a first shaft. A low pressure turbine is coupled to the low pressure compressor via a second shaft. A geared architecture interconnects between the second shaft and the fan. The gas turbine engine is a high bypass geared aircraft engine having a bypass ratio of greater than six (6). The low pressure turbine has a pressure ratio that is greater than 5, and the geared architecture includes a gear reduction ratio of greater than 2.5:1.

CROSS-REFERENCE TO RELATED APPLICATIONS

This disclosure is a continuation of U.S. patent application Ser. No.13/362,747 filed Jan. 31, 2012.

BACKGROUND

This disclosure relates to a gas turbine engine mid turbine framebearing support.

One typical gas turbine engine includes multiple, nested coaxial spools.A low pressure turbine is mounted on a first spool, and a high pressureturbine is mounted on a second spool. A mid turbine frame, which is partof the engine's static structure, is arranged axially between the lowand high pressure turbines. The turbine frame includes an inner hub andouter shroud with a circumferential array of airfoils adjoining the huband shroud, providing a gas flow path.

One typical static structure design provide a single, conical memberbetween “hot” components, like the gas flow path and supporting casestructures, and “cold” components, such as bearing compartments thatmust be kept at low temperatures to prevent oil coking. The conicalmember allows the cold and hot parts to axially shift relative to oneanother to accommodate the different thermal expansion rates of the coldand hot parts. This conical member is generally long in both the axialdirection, and in radial height when compared to the bearingcompartment.

SUMMARY

In one exemplary embodiment, a gas turbine engine includes high and lowpressure turbines. A mid-turbine frame is arranged axially between thehigh and low pressure turbines, and a bearing is operatively supportedby a support structure. An inner case is secured to the supportstructure and including a first conical member. A bearing support isprovided and to which the bearing is mounted. The bearing supportincludes a second conical member that is secured to the first conicalmember at a joint. The first and second conical members are arrangedradially inward of the joint. The joint is provided at a planeperpendicular to a rotational axis of the high and low pressureturbines. A first axial side is on one side of the plane, and a secondaxial side is on another side of the plane opposite the one side. Thefirst and second conical members are arranged on the first axial side. Afan is provided, and a compressor section is fluidly connected to thefan. The compressor includes a high pressure compressor and a lowpressure compressor. A combustor is fluidly connected to the compressorsection. A turbine section is fluidly connected to the combustor. Theturbine section includes the high pressure turbine coupled to the highpressure compressor via a first shaft. The low pressure turbine iscoupled to the low pressure compressor via a second shaft. A gearedarchitecture interconnects between the second shaft and the fan. The gasturbine engine is a high bypass geared aircraft engine having a bypassratio of greater than six (6).

In a further embodiment of the above, the low pressure turbine has apressure ratio that is greater than 5.

In a further embodiment of any of the above, the geared architectureincludes a gear reduction ratio of greater than 2.5:1.

In a further embodiment of any of the above, the low pressure turbinehas a pressure ratio that is greater than 5. The geared architectureincludes a gear reduction ratio of greater than 2.5:1.

In another exemplary embodiment, a gas turbine engine includes high andlow pressure turbines. A mid-turbine frame is arranged axially betweenthe high and low pressure turbines, and a bearing is operativelysupported by a support structure. An inner case is secured to thesupport structure and including a first conical member. A bearingsupport is provided and to which the bearing is mounted. The bearingsupport includes a second conical member that is secured to the firstconical member at a joint. The first and second conical members arearranged radially inward of the joint. The joint is provided at a planeperpendicular to a rotational axis of the high and low pressureturbines. A first axial side is on one side of the plane, and a secondaxial side is on another side of the plane opposite the one side. Thefirst and second conical members are arranged on the first axial side. Afan is provided, and a compressor section is fluidly connected to thefan. The compressor includes a high pressure compressor and a lowpressure compressor. A combustor is fluidly connected to the compressorsection. A turbine section is fluidly connected to the combustor. Theturbine section includes the high pressure turbine coupled to the highpressure compressor via a first shaft. The low pressure turbine iscoupled to the low pressure compressor via a second shaft. A gearedarchitecture interconnects between the second shaft and the fan. The lowpressure turbine has a pressure ratio that is greater than 5.

In a further embodiment of any of the above, the geared architectureincludes a gear reduction ratio of greater than 2.5:1.

In another exemplary embodiment, a gas turbine engine includes high andlow pressure turbines. A mid-turbine frame is arranged axially betweenthe high and low pressure turbines, and a bearing is operativelysupported by a support structure. An inner case is secured to thesupport structure and including a first conical member. A bearingsupport is provided and to which the bearing is mounted. The bearingsupport includes a second conical member that is secured to the firstconical member at a joint. The first and second conical members arearranged radially inward of the joint. The joint is provided at a planeperpendicular to a rotational axis of the high and low pressureturbines. A first axial side is on one side of the plane, and a secondaxial side is on another side of the plane opposite the one side. Thefirst and second conical members are arranged on the first axial side. Afan is provided, and a compressor section is fluidly connected to thefan. The compressor includes a high pressure compressor and a lowpressure compressor. A combustor is fluidly connected to the compressorsection. A turbine section is fluidly connected to the combustor. Theturbine section includes the high pressure turbine that is coupled tothe high pressure compressor via a first shaft. The low pressure turbineis coupled to the low pressure compressor via a second shaft. A gearedarchitecture interconnects between the second shaft and the fan. Thegeared architecture includes a gear reduction ratio of greater than2.5:1.

BRIEF DESCRIPTION OF THE DRAWINGS

The disclosure can be further understood by reference to the followingdetailed description when considered in connection with the accompanyingdrawings wherein:

FIG. 1 schematically illustrates a gas turbine engine.

FIG. 2 is a cross-sectional view of a portion of an engine staticstructure in the area of a mid turbine frame.

FIG. 3 is another example cross-sectional view of a portion of an enginestatic structure in the area of a mid turbine frame.

DETAILED DESCRIPTION

FIG. 1 schematically illustrates a gas turbine engine 20. The gasturbine engine 20 is disclosed herein as a two-spool turbofan thatgenerally incorporates a fan section 22, a compressor section 24, acombustor section 26 and a turbine section 28. Alternative engines mightinclude an augmentor section (not shown) among other systems orfeatures. The fan section 22 drives air along a bypass flowpath whilethe compressor section 24 drives air along a core flowpath forcompression and communication into the combustor section 26 thenexpansion through the turbine section 28. Although depicted as aturbofan gas turbine engine in the disclosed non-limiting embodiment, itshould be understood that the concepts described herein are not limitedto use with turbofans as the teachings may be applied to other types ofturbine engines including three-spool architectures.

The engine 20 generally includes a low speed spool 30 and a high speedspool 32 mounted for rotation about an engine central longitudinal axisA relative to an engine static structure 36 via several bearing systems38. It should be understood that various bearing systems 38 at variouslocations may alternatively or additionally be provided.

The low speed spool 30 generally includes an inner shaft 40 thatinterconnects a fan 42, a low pressure compressor 44 and a low pressureturbine 46. The inner shaft 40 is connected to the fan 42 through ageared architecture 48 to drive the fan 42 at a lower speed than the lowspeed spool 30. The high speed spool 32 includes an outer shaft 50 thatinterconnects a high pressure compressor 52 and high pressure turbine54. A combustor 56 is arranged between the high pressure compressor 52and the high pressure turbine 54. A mid-turbine frame 57 of the enginestatic structure 36 is arranged generally between the high pressureturbine 54 and the low pressure turbine 46. The mid-turbine frame 57supports one or more bearing systems 38 in the turbine section 28. Theinner shaft 40 and the outer shaft 50 are concentric and rotate viabearing systems 38 about the engine central longitudinal axis A, whichis collinear with their longitudinal axes.

The core airflow is compressed by the low pressure compressor 44 thenthe high pressure compressor 52, mixed and burned with fuel in thecombustor 56, then expanded over the high pressure turbine 54 and lowpressure turbine 46. The mid-turbine frame 57 includes airfoils 59 whichare in the core airflow path. The turbines 46, 54 rotationally drive therespective low speed spool 30 and high speed spool 32 in response to theexpansion.

The engine 20 in one example a high-bypass geared aircraft engine. In afurther example, the engine 20 bypass ratio is greater than about six(6), with an example embodiment being greater than a ratio ofapproximately 10:1, the geared architecture 48 is an epicyclic geartrain, such as a planetary gear system or other gear system, with a gearreduction ratio of greater than about 2.3:1 and the low pressure turbine46 has a pressure ratio that is greater than about 5. In one disclosedembodiment, the engine 20 bypass ratio is greater than about 10:1, thefan diameter is significantly larger than that of the low pressurecompressor 44, and the low pressure turbine 46 has a pressure ratio thatis greater than about 5:1. Low pressure turbine 46 pressure ratio ispressure measured prior to inlet of low pressure turbine 46 as relatedto the pressure at the outlet of the low pressure turbine 46 prior to anexhaust nozzle. The geared architecture 48 may be an epicycle geartrain, such as a planetary gear system or other gear system, with a gearreduction ratio of greater than about 2.5:1. It should be understood,however, that the above parameters are only exemplary of one embodimentof a geared architecture engine is applicable to other gas turbineengines including direct drive turbofans.

A significant amount of thrust is provided by the bypass flow B due tothe high bypass ratio. The fan section 22 of the engine 20 is designedfor a particular flight condition—typically cruise at about 0.8 Mach andabout 35,000 feet. The flight condition of 0.8 Mach and 35,000 ft, withthe engine at its best fuel consumption—also known as “bucket cruiseThrust Specific Fuel Consumption (‘TSFC’)”—is the industry standardparameter of lbm of fuel being burned divided by lbf of thrust theengine produces at that minimum point.

To make an accurate comparison of fuel consumption between engines, fuelconsumption is reduced to a common denominator, which is applicable toall types and sizes of turbojets and turbofans. The term is thrustspecific fuel consumption, or TSFC. This is an engine's fuel consumptionin pounds per hour divided by the net thrust. The result is the amountof fuel required to produce one pound of thrust. The TSFC unit is poundsper hour per pounds of thrust (lb/hr/lb Fn). When it is obvious that thereference is to a turbojet or turbofan engine, TSFC is often simplycalled specific fuel consumption, or SFC.

“Low fan pressure ratio” is the pressure ratio across the fan bladealone, without a Fan Exit Guide Vane (“FEGV”) system. The low fanpressure ratio as disclosed herein according to one non-limitingembodiment is less than about 1.45. “Low corrected fan tip speed” is theactual fan tip speed in ft/sec divided by an industry standardtemperature correction of [(Tambient deg R)/518.7)^0.5]. The “Lowcorrected fan tip speed” as disclosed herein according to onenon-limiting embodiment is less than about 1150 ft/second.

Referring to FIG. 2, the mid turbine frame 57 includes a hub 58supporting the airfoil 59. In one example, the mid turbine frame 57 is anickel alloy, but is not intended to carry the structural load of thebearing system 38 and its supported components. Instead, an inner case60 provided the support structure of the mid turbine frame 57 and issupported by radially extending, circumferentially arranged rods 64secured to the inner case 60 by a nut 66. The rods 64 are mounted tosupport structure 63.

The mid turbine frame 57 is a “hot” component that is isolated from thebearing system 38, a “cold” component. To this end, an air seal 68 isprovided to create a cooling cavity 86 between the inner case 60 and themid turbine frame 57. A cooling source 88, such as low compressorturbine air, is in fluid communication with the cooling cavity 86, forexample, through passages 62 provided in the airfoils 59.

A bearing support 90 is secured to the inner case 60 at a joint 96 withfasteners 98. A bearing 100 supports the outer shaft 50, for example,for rotation relative to the bearing support 90. The joint 96 isprovided structurally between the bearing 100 and the support structure63 (and rod 64 in the example). The bearing 100 is arranged in a bearingcompartment 102 sealed by a bearing compartment seal.

Instead of a single conical member, at least two conical members 92, 94,which may be nickel alloys, are used to provide structural support fromthe rods 64, in the example, to the bearing 100. The first and secondconical members 92, 94 connect the bearing 100 to the mid turbine frame57. Such an arrangement provides a more radially compact supportconfiguration while maintaining flexibility between the “hot” and “cold”components throughout various thermal gradients. The first and secondconical members 92, 94, or cones, are nested relative to one another andarranged radially inward of the joint 96. In the example shown in FIG.2, the first and second conical members 92, 94 are arranged on the sameaxial side of the joint 96. The first conical member 92 at leastpartially surrounds the second conical member 94.

In an example shown in FIG. 3, the first and second conical members 92,194 are arranged on opposing axial sides of the joint 196. The bearingsupport 190 is provided by an intermediate member 193 that supports thebearing 100 and is secured to the second conical member 194 at anintermediate joint 199.

Although an example embodiment has been disclosed, a worker of ordinaryskill in this art would recognize that certain modifications would comewithin the scope of the claims. For that reason, the following claimsshould be studied to determine their true scope and content.

What is claimed is:
 1. A gas turbine engine comprising: high and low pressure turbines; a mid-turbine frame arranged axially between the high and low pressure turbines; a bearing operatively supported by a support structure; an inner case secured to the support structure and including a first conical member; and a bearing support to which the bearing is mounted, the bearing support including a second conical member that is secured to the first conical member at a joint, the first and second conical members arranged radially inward of the joint, the joint provided at a plane perpendicular to a rotational axis of the high and low pressure turbines, a first axial side on one side of the plane, and a second axial side on another side of the plane opposite the one side, the first and second conical members are arranged on the first axial side; a fan; a compressor section fluidly connected to the fan, the compressor comprising a high pressure compressor and a low pressure compressor; a combustor fluidly connected to the compressor section; a turbine section fluidly connected to the combustor, the turbine section comprising: a high pressure turbine coupled to the high pressure compressor via a first shaft; a low pressure turbine coupled to the low pressure compressor via a second shaft; a geared architecture interconnects between the second shaft and the fan; and wherein the gas turbine engine is a high bypass geared aircraft engine having a bypass ratio of greater than six (6).
 2. The gas turbine engine according to claim 1, wherein the low pressure turbine has a pressure ratio that is greater than
 5. 3. The gas turbine engine according to claim 1, wherein the geared architecture includes a gear reduction ratio of greater than 2.5:1.
 4. The gas turbine engine according to claim 1, wherein the low pressure turbine has a pressure ratio that is greater than 5, and wherein the geared architecture includes a gear reduction ratio of greater than 2.5:1.
 5. A gas turbine engine comprising: high and low pressure turbines; a mid-turbine frame arranged axially between the high and low pressure turbines; a bearing operatively supported by a support structure; an inner case secured to the support structure and including a first conical member; and a bearing support to which the bearing is mounted, the bearing support including a second conical member that is secured to the first conical member at a joint, the first and second conical members arranged radially inward of the joint, the joint provided at a plane perpendicular to a rotational axis of the high and low pressure turbines, a first axial side on one side of the plane, and a second axial side on another side of the plane opposite the one side, the first and second conical members are arranged on the first axial side; a fan; a compressor section fluidly connected to the fan, the compressor comprising a high pressure compressor and a low pressure compressor; a combustor fluidly connected to the compressor section; a turbine section fluidly connected to the combustor, the turbine section comprising: a high pressure turbine coupled to the high pressure compressor via a first shaft; a low pressure turbine coupled to the low pressure compressor via a second shaft; a geared architecture interconnects between the second shaft and the fan; and wherein the low pressure turbine has a pressure ratio that is greater than
 5. 6. The gas turbine engine according to claim 5, wherein the geared architecture includes a gear reduction ratio of greater than 2.5:1.
 7. A gas turbine engine comprising: high and low pressure turbines; a mid-turbine frame arranged axially between the high and low pressure turbines; a bearing operatively supported by a support structure; an inner case secured to the support structure and including a first conical member; and a bearing support to which the bearing is mounted, the bearing support including a second conical member that is secured to the first conical member at a joint, the first and second conical members arranged radially inward of the joint, the joint provided at a plane perpendicular to a rotational axis of the high and low pressure turbines, a first axial side on one side of the plane, and a second axial side on another side of the plane opposite the one side, the first and second conical members are arranged on the first axial side; a fan; a compressor section fluidly connected to the fan, the compressor comprising a high pressure compressor and a low pressure compressor; a combustor fluidly connected to the compressor section; a turbine section fluidly connected to the combustor, the turbine section comprising: a high pressure turbine coupled to the high pressure compressor via a first shaft; a low pressure turbine coupled to the low pressure compressor via a second shaft; a geared architecture interconnects between the second shaft and the fan; and wherein the geared architecture includes a gear reduction ratio of greater than 2.5:1. 